Flyaway stringer end caps

ABSTRACT

Systems and methods are provided for fabricating composite parts. One embodiment is a method for fabricating a composite part. The method includes forming a skin panel preform comprising fiber reinforced material, disposing rigid end caps at the skin panel preform at end locations of stringer preforms that will be placed at the skin panel, locating the stringer preforms at the skin panel preform via the rigid end caps, and anchoring the stringer preforms to the skin panel preform.

FIELD

The disclosure relates to the field of fabrication, and in particular,to fabrication of stringers for aircraft.

BACKGROUND

Stringers for an aircraft (e.g., the wings of an aircraft) may befabricated from composite materials. These stringers may be laid-up aspreforms, processed into stringers, and then co-cured to a wing skin inorder to form a completed wing or portion thereof. When stringers arefabricated in this manner, the stringers are laid-up and cured such thatthey terminate within the wing skin. The stringers can then be cut todesired dimensions at desired ramp rates after curing has completed. Thetransition from stringer to skin is commonly referred to as a “stringerrun-out” and is utilized for transferring load in composite wingstructures.

Fabricating a stringer made of composite materials in this manner mayresult in edge conditions which are out of tolerance or, and out oftolerance conditions are undesirable as they may necessitate rework. Atthe same time, it may be difficult to lay up and cure a preform atdesired ramp rates, owing to the complex geometries that may berequired.

Therefore, it would be desirable to have a method and apparatus thattake into account at least some of the issues discussed above, as wellas other possible issues.

SUMMARY

Embodiments described herein provide rigid end caps for stringers thatexhibit desired ramp rates and stringer run-out quality for terminatinga stringer. These end caps are affixed to preforms for the stringers toform integral portions of a composite part. The end caps eliminate theneed to shape the preforms to specified ramp rates, while alsoeliminating the need to cut the stringers after curing. The end capsform flyaway tooling that enforces desired shapes onto stringers duringcuring, while also bearing and transferring loads within the wing afterthe wing has been fabricated. One embodiment is a method of forming astringer. The method includes laying up a stringer preform comprisingfiber-reinforced material, placing the stringer preform onto a skinpanel preform, bonding an end cap to the stringer preform and the skinpanel preform, and co-curing the stringer preform and the skin panelpreform while the end cap is bonded to the stringer preform and the skinpanel preform, resulting in a composite part that includes the end cap.

One embodiment is a method for fabricating a composite part. The methodincludes forming a skin panel preform comprising fiber reinforcedmaterial, disposing rigid end caps at the skin panel preform at endlocations of stringer preforms that will be placed at the skin panel,locating the stringer preforms at the skin panel preform via the rigidend caps, and anchoring the stringer preforms to the skin panel preform.

A further embodiment is an apparatus for receiving a stringer preform.The apparatus includes a skin panel preform comprising fiber reinforcedmaterial, and rigid end caps that are disposed atop the skin panelpreform and are separated by a length of a stringer preform that will beplaced onto the skin panel preform.

A further embodiment is a non-transitory computer readable mediumembodying programmed instructions which, when executed by a processor,are operable for performing a method for forming a stringer. The methodincludes forming a skin panel preform comprising fiber reinforcedmaterial, disposing rigid end caps at the skin panel preform at endlocations of stringer preforms that will be placed at the skin panel,locating the stringer preforms at the skin panel preform via the rigidend caps, and anchoring the stringer preforms to the skin panel preform.

A further embodiment is a system that forms a portion of an aircraft.The system includes a section of airframe comprising a skin panelcomprising fiber reinforced material, stringers that are affixed to theskin panel and that comprise fiber reinforced material, and rigid endcaps that are integral with the skin panel and the stringers, the rigidend caps comprising flyaway tooling that supported the stringers duringhardening.

Other illustrative embodiments (e.g., methods and computer-readablemedia relating to the foregoing embodiments) may be described below. Thefeatures, functions, and advantages that have been discussed can beachieved independently in various embodiments or may be combined in yetother embodiments further details of which can be seen with reference tothe following description and drawings.

DESCRIPTION OF THE DRAWINGS

Some embodiments of the present disclosure are now described, by way ofexample only, and with reference to the accompanying drawings. The samereference number represents the same element or the same type of elementon all drawings.

FIG. 1 is a block diagram of one half of a wing that includes co-curedstringers and flyaway tooling in an illustrative embodiment.

FIG. 2A is a flowchart illustrating a method for integrating flyawaytooling into a wing in an illustrative embodiment.

FIG. 2B is a flowchart illustrating a method for integrating flyawaytooling into a wing in an illustrative embodiment.

FIGS. 3 and 4A-4C illustrate an end cap for a first stringer shape in anillustrative embodiment.

FIGS. 5-6 illustrate an end cap for a second stringer shape in anillustrative embodiment.

FIGS. 7A-7E and 8 illustrate a co-bonded end cap and stringer in anillustrative embodiment.

FIG. 9 is a perspective view of a wing that includes flyaway tooling inan illustrative embodiment.

FIG. 10 illustrates an inboard end cap for a stringer in an illustrativeembodiment.

FIG. 11 is a block diagram illustrating inboard end caps for stringersintegrated into an airframe in an illustrative embodiment.

FIG. 12 is a further flowchart illustrating a method for integratingflyaway tooling into a wing in an illustrative embodiment.

FIG. 13 is a flow diagram of aircraft production and service methodologyin an illustrative embodiment.

FIG. 14 is a block diagram of an aircraft in an illustrative embodiment.

DESCRIPTION

The figures and the following description provide specific illustrativeembodiments of the disclosure. It will thus be appreciated that thoseskilled in the art will be able to devise various arrangements that,although not explicitly described or shown herein, embody the principlesof the disclosure and are included within the scope of the disclosure.Furthermore, any examples described herein are intended to aid inunderstanding the principles of the disclosure, and are to be construedas being without limitation to such specifically recited examples andconditions. As a result, the disclosure is not limited to the specificembodiments or examples described below, but by the claims and theirequivalents.

Composite parts, such as Carbon Fiber Reinforced Polymer (CFRP) parts,are initially laid-up in multiple layers that together are referred toas a preform. Individual fibers within each layer of the preform arealigned parallel with each other, but different layers exhibit differentfiber orientations in order to increase the strength of the resultingcomposite part along different dimensions. The preform includes aviscous resin that solidifies in order to harden the preform into acomposite part (e.g., for use in an aircraft). Carbon fiber that hasbeen impregnated with an uncured thermoset resin or a thermoplasticresin is referred to as “prepreg.” Other types of carbon fiber include“dry fiber” which has not been impregnated with thermoset resin but mayinclude a tackifier or binder. Dry fiber is infused with resin prior tocuring. For thermoset resins, the hardening is a one-way processreferred to as curing, while for thermoplastic resins, the resin reachesa viscous form if it is re-heated.

FIG. 1 is a block diagram of a lower half of a wing 100 that includesco-cured stringers and flyaway tooling in an illustrative embodiment.The lower half of the wing 100 illustrated in FIG. 1 may comprise aco-cured lower half of a wing, and is capable of being affixed or bondedto an upper half of a wing in order to form a complete wing. In furtherembodiments, an upper half of a wing is assembled in a similar manner tothat of the lower half of the wing 100. The lower half of a wing 100depicted in FIG. 1 includes fiber-reinforced composite materials as wellas rigid end caps for stringers. This arrangement eliminates the needfor trimming or cutting the stringers at the wing 100 after hardening.

In this embodiment, FIG. 1 depicts a portion of a wing that includes askin panel 110, which comprises fiber-reinforced material in the form ofmultiple layers 112 of resin 114 and fibers 116. A stringer 120 isdisposed atop the skin panel 110, and comprises fiber-reinforcedmaterial in the form of layers 112 of fibers 116 and resin 114. Thestringer 120 is co-cured to the skin panel 110.

End caps 130 abut the inboard end 122 and outboard end 124 of thestringer 120. The end caps 130 are rigid prior to curing, and maycomprise a metal that provides high strength with low weight, such astitanium or aluminum (e.g., isolated from carbon fiber by isolationplies of fiberglass or other material). End caps 130 may furthercomprise hardened composite materials (e.g., thermoset orthermoplastic), as well as 3D-printed metals. Thus, in one embodimentthe end caps 130 are fabricated via additive manufacturing techniques,such as 3D printing. In further embodiments, subtractive manufacturingtechniques are utilized. Each end cap 130 includes a flange 132 with aramp 134 for receiving the stringer 120. The end caps 130 are boltedand/or bonded to the stringer 120 and/or skin panel 110.

The ramp 134 provides a pathway for transferring load from the stringer120 to the end cap 130. As used herein, a “ramp” refers to any physicalstructure that transitions load along its length, including step laps,scarfing, interleaving, linear ramps, and other features. The stringer120 itself includes a flange and web, and these structures arecomplementary to the ramp 134 of the end cap 130. Each end cap 130 alsoincludes a ramp 136 that proceeds down to the skin panel 110. The ramp136 provides a pathway for transferring load from the end cap 130 to theskin panel 110.

The end caps 130 are co-bonded to the stringers 120 and the skin panel110, and the ramps 134 of the end caps 130 are overlapped withfiber-reinforced material from the stringers 120 in one embodiment. Infurther embodiments, the ramps 134 and/or 136 integrate with stringers120 via laps, step laps, or scarf interleaving of the ramps with theplies of the stringer and/or skin panel 110. In some of theseembodiments, the transitions involve laying up a stinger preform upon anend cap 130 to accommodate differences in shape. The terms of “skinpanel 110” and “stringer 120” are utilized herein to refer to bothuncured preforms as well as hardened composite parts. That is, a skinpanel 110 may refer to an unhardened preform for a skin panel awaitingcuring, or may refer to a hardened skin panel. In a similar fashion, astringer 120 may refer to an unhardened preform for a stringer, or to ahardened stringer.

FIG. 1 further depicts components which may be electronically managed bycontroller 140 to fabricate the structure discussed above. In thisembodiment, the components include a layup machine 142, such as anAutomated Fiber Placement (AFP) machine or tape dispensing head thatlays up tows of unidirectional fiber-reinforced material to form thelayers 112. The components further include a Pick and Place (PNP)machine 144 (e.g., an end effector, suction device, gripper, etc.),which picks up and places stringer preforms and/or end caps 130 ontoskin panel 110. In further embodiments, the stringer preforms and endcaps 130 are manually picked and placed into position by one or moretechnicians. Controller 140 directs the operations of these componentsbased on instructions stored in one or more Numerical Control (NC)programs in memory, and may be implemented, for example, as customcircuitry, as a hardware processor executing programmed instructions, orsome combination thereof.

Illustrative details of the operation of the components of FIG. 1 willbe discussed with regard to FIG. 2A. Assume, for this embodiment, that askin panel preform (e.g., skin panel 110) for a wing has been laid-up bylayup machine 142.

FIG. 2A is a flowchart illustrating a method 200 for integrating flyawaytooling into a wing in an illustrative embodiment. The steps of method200 are described with reference to the lower half of a wing 100 shownin FIG. 1 , but those skilled in the art will appreciate that method 200may be performed for other portions of wings (e.g., upper halves ofwings) in other environments. The steps of the flowcharts describedherein are not all inclusive and may include other steps not shown. Thesteps described herein may also be performed in an alternative order.

In step 202, controller 140 directs layup machine 142 to lay up thestringer preform. In one embodiment, this comprises applying multiplelayers of unidirectional fiber-reinforced material to a layup mandrel(not shown) or other piece of tooling that defines a shape for thestringer preform.

In one embodiment, the PNP machine 144 places an end cap 130 at each endof the stringer preform. The end caps 130 help to enforce a desiredshape at the stringer preform before and during curing. Furthermore, theend caps 130 provide ramps 136 (e.g., for runouts) and/or other complexgeometries in a rigid form, which means that these geometries do notneed to be mechanically supported during the curing process. Thisreduces the complexity of layup and curing for the wing, which reducesexpenses related to labor and materials.

In step 204, controller 140 directs the PNP machine 144 to place thestringer preform onto a skin panel preform (e.g., skin panel 110).During this operation, the stringer preform has not yet been cured(i.e., is still in the “green state”) and therefore remains flaccid. ThePNP machine 144 may therefore enforce or retain a desired curvature atthe stringer preform via the application of suction (e.g., via a vacuumconnection) or use of supporting structure while the stringer preform isbeing transported. In one embodiment, the PNP machine 144 picks up andplaces multiple stringer preforms at the skin panel preform. This mayalso include placing a stringer preform onto a layup mandrel and thenlaying up the skin against the stringer preform. For a skin panel thatdefines an upper portion of a wing, the stringers may comprise preformsfor hat stringers. For a skin panel that defines a lower portion of awing, the stringers may comprise preforms for “T” stringers. Furthertypes of stringers include Z stringers, and stringers of any suitablecross-section. In one embodiment, the placement operation involvesoverlapping a ramp 134 of the end cap 130 with one or more layers of thestringer preform. This may comprise overlapping the ramp 134 with a rampat the stringer preform. In this manner, after the stringer preformhardens, the ramp 134 transfers loads between the hardened stringer andthe end cap 130. Meanwhile the ramp 136 transfers loads between the endcap 130 and the skin panel 110.

Step 206 comprises bonding an end cap 130 to the stringer preform andthe skin panel preform. In one embodiment, this comprises applying anadhesive (e.g., an epoxy, glue, or other self-hardening chemical) to theend caps 130 prior to placing the end caps or the stringer preforms ontothe skin panel preform, and waiting for the adhesive to harden afterplacing the end caps and stringer preform into position. In oneembodiment, an end cap 130 is bonded to an outboard end of the stringerpreform, and another end cap 130 is bonded to an inboard end of thestringer preform.

After the end caps 130 have been bonded into place, the position of endcaps 130 with regard to the stringer preform and the skin panel preformis held in place by the hardened adhesive. This ensures that vacuumbagging setup and consolidation will not shift the position of thestringer preforms, skin panel preform, and end caps with respect to eachother. Thus, PNP machine 144, another machine, or a technician, mayproceed to vacuum bag the end cap, stringer preform, and skin panelpreform (i.e., prior to co-curing these elements together). The end capsboth anchor the stringer preforms to the skin panel preform, and locatethe stringers at the skin panel preform. In one embodiment, the vacuumbag is utilized to consolidate these components via the application ofpressure, prior to curing.

Step 208 comprises co-curing the stringer preform to the skin panelpreform while the end caps 130 are bonded to the stringer preform andthe skin panel preform, resulting in a composite part that includes theend caps 130 as integral components. In one embodiment, co-curingcomprises placing the vacuum-bagged components into an autoclave,applying heat via the autoclave until resin 114 reaches a curingtemperature, and applying pressure via the vacuum bag and/or via theautoclave in order to consolidate and cure the end caps 130, stringerpreforms, and skin panel preform. These components are co-cured into anintegral composite part that includes integral flyaway tooling. That is,the end caps 130 operate as tooling to provide support for the preformsduring curing and vacuum bagging, and also provide mechanical strengthwhen the resulting portion of wing is assembled into a portion of anairframe of an aircraft. The tooling becomes physically integral withstringers after hardening. Therefore, the tooling is flyaway toolingbecause it is integrated into an aircraft and “flies away” as part ofthe aircraft after fabrication has completed.

Method 200 provides a technical benefit over prior techniques, becauseit enables a runout to be rapidly integrated into a composite stringer,provides support during vacuum bagging and curing. This is because theend caps, being rigid prior to curing, resist compaction forces appliedby a vacuum bag that could crush or bend elongated portions such as aweb of a stringer preform. Hence, the end caps help to constrainstringer preforms to desired shapes during curing. Method 200additionally eliminates the need to cut or remove material from acomposite stringer after the composite stringer has been co-cured to askin panel.

FIG. 2B illustrates a method 250 for integrating flyaway tooling into awing in an illustrative embodiment. Method 250 comprises forming (e.g.,laying up) a skin panel preform comprising fiber reinforced material(e.g., CFRP) in step 252, disposing rigid end caps at the skin panelpreform at end locations of stringer preforms (e.g., laid-up from CFRP)that will be placed at the skin panel in step 254, locating the stringerpreforms at the skin panel preform via the rigid end caps in step 256,and anchoring the stringer preforms to the skin panel preform in step258.

In a further embodiment, method 250 includes consolidating the stringerpreforms and the skin panel preform via a vacuum bag that covers thestringer preforms, skin panel preform, and end caps. Method 250 mayfurther comprise hardening the stringer preforms and the skin panelpreform to form a section of wing that includes the end caps. The endcaps are bonded to the stringer preforms and the skin panel preform, andin one embodiment the method 250 further comprises co-curing thestringer preform and the skin panel preform while the rigid end caps arebonded to the stringer preform and the skin panel preform, resulting ina composite part that includes the end cap.

FIGS. 3 and 4A-4C illustrate an end cap 300 for a first stringer shapein an illustrative embodiment. FIG. 4A corresponds with view arrows 4Aof FIG. 3 . In this embodiment, the end cap 300 is designed as an endcap for a stringer having a T-shaped cross-section. That is, the end cap300 has a T-shaped cross-section that aligns with a T-shaped crosssection of a stringer.

The end cap 300 includes a web 310 which narrows via ramp 312, whichextends from the web 310 and tapers the web 310. In this embodiment, theweb 310 forms a vertical plane. However, in further embodiments the webforms a curved shape (e.g., as shown in FIG. 5 ). The web 310 protrudesfrom a lower flange 320. Lower flange 320 narrows via ramp 322, whichextends from the lower flange 320 and tapers the lower flange 320. Theseramps 312 and 322 are overlapped by corresponding ramps in the stringerpreform, such that an overall thickness of the combination of flange andstringer preform remains constant along the flange. In furtherembodiments, the ramps exhibit a stairstep pattern that accommodates astep lap or other type of interface/transition between the stringer andthe end cap. In still further embodiments, interleaving of compositeplies with metallic structure is performed to narrow the structureinstead of narrowing in a linear ramped fashion. Ramp 330 terminates thestringer in accordance with desired structural constraints, and isshaped and tapered to help transition the stringer load through the endcap to the panel. Ramp 340 transfers forces through the end cap 300 intoa skin panel (not shown). Meanwhile, the ramp 312 proceeding into theweb 310, and the ramp 322 proceeding into the lower flange 320 transferload from the stringer. The rate of transition of load through the endcap is established by the geometry and of the ramps and the pattern(e.g., linear, step, etc.) of the ramps.

FIG. 4B depicts a section-cut side view of the end cap 300 of FIG. 3 ,after a stringer 400 has been bonded thereto. FIG. 4B corresponds withview arrows 4B of FIG. 4A. In FIG. 4B, it can be seen that stringer 400includes ramp 410 and ramp 420, which mate with ramps 312 and 322,respectively. FIG. 4C is a section cut end view that corresponds withview arrows 4C of FIG. 4A. In FIG. 4C ramps 312 and 322 of end cap 300are visible, as are ramps 410 and 420 of stringer 400. In furtherembodiments the end cap is bonded to an outboard end of the stringerpreform and fastened to the wing panel, and another end cap is bonded toan inboard end of the stringer preform and fastened to the wing panel.

FIGS. 5-6 illustrate an end cap 500 for a second stringer shape in anillustrative embodiment. FIG. 6 corresponds with view arrows 6 of FIG. 5. In this embodiment, the end cap 500 is designed as an end cap for astringer having a “hat” shaped cross-section. That is, the end cap 500has a hat-shaped cross-section that aligns with a hat-shaped crosssection of a stringer.

The end cap 500 includes an upper arch 510 which narrows via ramp 512,and also includes a lower flange 520, which narrows via ramp 522. Theseflanges are overlapped by corresponding ramps in the stringer preform,such that an overall thickness of the combination of flange and stringerpreform remains constant along the flange. In further embodiments, theramps exhibit a stairstep pattern as they narrow, instead of narrowingin a linear fashion. Ramp 530 terminates the stringer in accordance withdesired structural constraints, and lip 540 distributes forces bornethrough the end cap 500 into a skin panel (not shown). In thisembodiment, the upper arch 510 forms a void 550 which corresponds with avoid in the hat stringer.

FIGS. 7A-7E and 8 illustrate an end cap 500 that has been co-bonded to astringer 700 in an illustrative embodiment. FIG. 8 is a side viewcorresponding with view arrows 8 of FIG. 7A. In this embodiment, ramp512 and ramp 522 are covered by fiber-reinforced material from thestringer 700, which forms a corresponding instance of a ramp 710 tomaintain a desired combined thickness of the stringer and end cap.

FIG. 7B is a section-cut side view that corresponds with view arrows 7Bof FIG. 7A, and depicts the ramp 710 intersecting with ramp 522 of FIG.7A at scarf joint 720. In FIG. 7B, the end cap has been bonded to a rampof a stringer 700, which structurally unites the stringer 700 with theend cap 500, forming a single component made from the stringer as wellas the end cap. Specifically, FIG. 7B depicts a scarf type of joinbetween the ramp of the end cap and the composite plies. However, a steptype of join can also be implemented, as depicted by step joint 730 ofFIG. 7C. In further embodiments, the end cap is interleaved withcomposite plies, as shown by interleaved joint 740 of FIG. 7D. Stillfurther implementations of interleaving may also or alternatively beutilized.

FIG. 7E is a section-cut end view that corresponds with view arrows 7Eof FIG. 7A. This view further depicts and intersection between ramps foran end cap and ramps for a stringer. As shown in FIG. 7E, a portion ofthe stringer 700 overlaps with the ramp 512 and ramp 522 of the end cap500.

FIG. 8 depicts a side view that corresponds with view arrows 8 of FIG.7A. In FIG. 8 , it can be observed that a majority of the join (i.e.,proceeding into the page between the ramp 710 and the ramp 522) ishidden from view.

FIG. 9 is a perspective view of a wing 900 that includes flyaway toolingin an illustrative embodiment. The tooling comprises end caps thatfacilitated hardening of the stringers by supporting the stringersduring curing, and that became physically integral with the stringersafter hardening. Therefore, the tooling is flyaway tooling because it isintegrated into an aircraft and “flies away” as part of the aircraftafter fabrication has completed. Thus, each end cap plays a dual role astooling that supports preforms prior to and during curing, and also as acomponent that transfers load from the stringer into the panel whenintegrated into an aircraft.

FIG. 9 illustrates that the wing 900 includes leading edge 902, trailingedge 904, and pylon 906. Wing 900 further includes stringers 920, aswell as inboard end caps 910 and outboard end caps 912 disposed at theinboard and outboard ends of the stringers 920. For example, the inboardend caps 910 may terminate a stringer 920 at a side-of-body intersectionto form an inboard end of a stringer 920, and the outboard end caps 912may terminate an outboard end of a stringer 920. In one embodiment,outboard end caps are smaller than inboard end caps, because they matewith tapered portions of the stringers 920. In such an embodimentinboard end caps do not taper into a skin panel, but rather are bondedor fastened to a wing box. Such end caps still taper into the stringersthemselves, however. Because FIG. 9 provides a simplified view, it willbe understood that there can be many more stringers than illustrated,and stringers can be much narrower relative to wing size. For example, astringer can taper from inboard end to outboard end. Also, some of thestringers may continue to the tip of the wing while others terminateearlier.

FIG. 10 illustrates an inboard end cap 1000 for a stringer in anillustrative embodiment. In this embodiment, the inboard end cap 1000 isdesigned as an end cap for a stringer having a T-shaped cross-section.That is, the end cap 1000 has a T-shaped cross-section that aligns witha T-shaped cross section of an inboard end of a stringer.

The inboard end cap 1000 includes a web 1010 which narrows via ramp1012, which extends from the web 1010 and tapers the web 1010. In thisembodiment, the web 1010 forms a vertical plane. The web 1010 protrudesfrom a lower flange 1020. Lower flange 1020 narrows via ramp 1022, whichextends from the lower flange 1020 and tapers the lower flange 1020.These ramps 1012 and 1022 are overlapped by/spliced with correspondingramps in the stringer preform, such that an overall thickness of thecombination of flange and stringer preform remains constant along theflange at end 1080. In further embodiments, the ramps exhibit astairstep pattern that accommodates a step lap or other type ofinterface/transition between the stringer and the end cap. In stillfurther embodiments, interleaving of composite plies with metallicstructure is performed to narrow the structure instead of narrowing in alinear ramped fashion. Web extension 1030 terminates the stringer byabutting against a side of body intersection at end 1080 to transferload from a wing box. Meanwhile, the ramp 1012 proceeding into the web1010, and the ramp 1022 proceeding into the lower flange 1020 transferload from the stringer to a center of the wing box. The rate oftransition of load through the end cap is established by the geometryand of the ramps and the pattern (e.g., linear, step, etc.) of the rampsand web extension.

FIG. 10 further depicts mounting features 1050 in the web and mountingfeature 1060 in the flange at the inboard end cap 1000. These mountingfeatures (e.g., bolt holes for fasteners, protrusions, etc.) facilitatethe process of affixing the inboard end cap 1000 to a center wing box.

FIG. 11 is a block diagram illustrating inboard end caps for stringersintegrated into an airframe in an illustrative embodiment. In FIG. 11 ,a side of body 1110 of an airframe is abutted by inboard end caps 1112.Stringers 1120 are attached to the inboard end caps 1112, and are alsoattached to outboard end caps 1114. The stringers 1120 and end caps arealso affixed to a skin panel 1104.

FIG. 12 is a further flowchart illustrating a method for integratingflyaway tooling into a wing in an illustrative embodiment. Step 1202comprises forming a skin panel preform via layup or other techniques.Step 1204 comprises bonding a caul tooling end (e.g., an end cap) ontothe skin panel preform. Step 1206 comprises anchoring a stringer preformat the skin panel preform and the caul tooling end. Step 1208 compriseshardening the skin panel preform and the stringer preform with the caultooling end. That is, while the skin panel preform and the stringerpreform are supported by the caul tooling end, they are hardened into acomposite part. In a further embodiment, vacuum bagging is performedprior to hardening, and this process includes vacuum bagging thestringer preform, the caul tooling end, and the skin panel preform. Inone embodiment, hardening comprises curing the stringer preform, thecaul tooling end, and the skin panel preform into a composite part whileapplying pressure. In a further embodiment, the caul tooling end formsflyaway tooling that structurally supports ends of a stringer. In afurther embodiment, the caul tooling end comprises a material selectedfrom the group consisting of metal and Carbon Fiber Reinforced Polymer(CFRP).

EXAMPLES

In the following examples, additional processes, systems, and methodsare described in the context of end caps used as flyaway tooling forstringers.

Referring more particularly to the drawings, embodiments of thedisclosure may be described in the context of aircraft manufacturing andservice in method 1200 as shown in FIG. 12 and an aircraft 1202 as shownin FIG. 13 . During pre-production, method 1200 may includespecification and design 1204 of the aircraft 1202 and materialprocurement 1206. During production, component and subassemblymanufacturing 1208 and system integration 1210 of the aircraft 1202takes place. Thereafter, the aircraft 1202 may go through certificationand delivery 1212 in order to be placed in service 1214. While inservice by a customer, the aircraft 1202 is scheduled for routine workin maintenance and service 1216 (which may also include modification,reconfiguration, refurbishment, and so on). Apparatus and methodsembodied herein may be employed during any one or more suitable stagesof the production and service described in method 1200 (e.g.,specification and design 1204, material procurement 1206, component andsubassembly manufacturing 1208, system integration 1210, certificationand delivery 1212, service 1214, maintenance and service 1216) and/orany suitable component of aircraft 1202 (e.g., airframe 1218, systems1220, interior 1222, propulsion system 1224, electrical system 1226,hydraulic system 1228, environmental 1230).

Each of the processes of method 1200 may be performed or carried out bya system integrator, a third party, and/or an operator (e.g., acustomer). For the purposes of this description, a system integrator mayinclude without limitation any number of aircraft manufacturers andmajor-system subcontractors; a third party may include withoutlimitation any number of vendors, subcontractors, and suppliers; and anoperator may be an airline, leasing company, military entity, serviceorganization, and so on.

As shown in FIG. 12 , the aircraft 1202 produced by method 1200 mayinclude an airframe 1218 with a plurality of systems 1220 and aninterior 1222. Examples of systems 1220 include one or more of apropulsion system 1224, an electrical system 1226, a hydraulic system1228, and an environmental system 1230. Any number of other systems maybe included. Although an aerospace example is shown, the principles ofthe invention may be applied to other industries, such as the automotiveindustry.

As already mentioned above, apparatus and methods embodied herein may beemployed during any one or more of the stages of the production andservice described in method 1200. For example, components orsubassemblies corresponding to component and subassembly manufacturing1208 may be fabricated or manufactured in a manner similar to componentsor subassemblies produced while the aircraft 1202 is in service. Also,one or more apparatus embodiments, method embodiments, or a combinationthereof may be utilized during the subassembly manufacturing 1208 andsystem integration 1212, for example, by substantially expeditingassembly of or reducing the cost of an aircraft 1202. Similarly, one ormore of apparatus embodiments, method embodiments, or a combinationthereof may be utilized while the aircraft 1202 is in service, forexample and without limitation during the maintenance and service 1216.Thus, the invention may be used in any stages discussed herein, or anycombination thereof, such as specification and design 1204, materialprocurement 1206, component and subassembly manufacturing 1208, systemintegration 1210, certification and delivery 1212, service 1214,maintenance and service 1216 and/or any suitable component of aircraft1202 (e.g., airframe 1218, systems 1220, interior 1222, propulsionsystem 1224, electrical system 1226, hydraulic system 1228, and/orenvironmental 1230.)

In one embodiment, a part comprises a portion of airframe 1218, and ismanufactured during component and subassembly manufacturing 1208. Thepart may then be assembled into an aircraft in system integration 1210,and then be utilized in service 1214 until wear renders the partunusable. Then, in maintenance and service 1216, the part may bediscarded and replaced with a newly manufactured part. Inventivecomponents and methods may be utilized throughout component andsubassembly manufacturing 1208 in order to manufacture new parts.

Any of the various control elements (e.g., electrical or electroniccomponents) shown in the figures or described herein may be implementedas hardware, a processor implementing software, a processor implementingfirmware, or some combination of these. For example, an element may beimplemented as dedicated hardware. Dedicated hardware elements may bereferred to as “processors”, “controllers”, or some similar terminology.When provided by a processor, the functions may be provided by a singlededicated processor, by a single shared processor, or by a plurality ofindividual processors, some of which may be shared. Moreover, explicituse of the term “processor” or “controller” should not be construed torefer exclusively to hardware capable of executing software, and mayimplicitly include, without limitation, digital signal processor (DSP)hardware, a network processor, application specific integrated circuit(ASIC) or other circuitry, field programmable gate array (FPGA), readonly memory (ROM) for storing software, random access memory (RAM),non-volatile storage, logic, or some other physical hardware componentor module.

Also, a control element may be implemented as instructions executable bya processor or a computer to perform the functions of the element. Someexamples of instructions are software, program code, and firmware. Theinstructions are operational when executed by the processor to directthe processor to perform the functions of the element. The instructionsmay be stored on storage devices that are readable by the processor.Some examples of the storage devices are digital or solid-statememories, magnetic storage media such as a magnetic disks and magnetictapes, hard drives, or optically readable digital data storage media.

Although specific embodiments are described herein, the scope of thedisclosure is not limited to those specific embodiments. The scope ofthe disclosure is defined by the following claims and any equivalentsthereof.

What is claimed is:
 1. A method for fabricating a composite part, themethod comprising: forming a skin panel preform comprising fiberreinforced material; disposing rigid end caps at the skin panel preformat end locations of stringer preforms that will be placed at the skinpanel preform; locating the stringer preforms at the skin panel preformvia the rigid end caps disposed at the skin panel preform; and anchoringthe stringer preforms to the skin panel preform between the rigid endcaps.
 2. The method of claim 1 further comprising: consolidating thestringer preforms and the skin panel preform via a vacuum bag thatcovers the stringer preforms, the skin panel preform, and the end caps.3. The method of claim 1 further comprising: hardening the stringerpreforms and the skin panel preform to form a section of wing thatincludes the end caps.
 4. The method of claim 1 further comprising:bonding the end caps to the stringer preforms.
 5. The method of claim 1further comprising: bonding the end caps to the skin panel preform. 6.The method of claim 1 further comprising: fastening the end caps to theskin panel preform.
 7. The method of claim 1 further comprising:co-curing the stringer preform and the skin panel preform while therigid end caps are bonded to the stringer preform and the skin panelpreform, resulting in a composite part that includes the end cap.
 8. Themethod of claim 1 further comprising: laying up the skin panel preformfrom Carbon Fiber Reinforced Polymer (CFRP).
 9. The method of claim 1further comprising: laying up the stringer preforms from Carbon FiberReinforced Polymer (CFRP).
 10. The method of claim 1 wherein: locating astringer preform comprises abutting the stringer preform against an endcap.
 11. A portion of an aircraft assembled according to the method ofclaim
 1. 12. The method of claim 1 wherein: locating the stringerpreforms at the skin panel preform comprises aligning cross-sections ofthe stringer preforms with cross-sections of the end caps.
 13. Themethod of claim 1 further comprising: directing a Pick and Place (PNP)machine to place the stringer preforms onto the skin panel preform. 14.The method of claim 13 further comprising: utilizing a Numerical Control(NC) program to direct the PNP machine.
 15. The method of claim 1wherein: locating the stringer preforms at the skin panel preformcomprises overlapping ramps of the end caps with one or more layers ofthe stringer preforms.
 16. The method of claim 1 further comprising:transferring loads between a ramp of a stringer preform and an end cap.17. The method of claim 1 further comprising: waiting for an adhesive toharden after locating the stringer preforms at the skin panel preform.18. The method of claim 1 further comprising: placing the stringerpreforms and the skin panel preform into an autoclave.
 19. The method ofclaim 1 further comprising: aligning the stringer preforms withstairstep patterns at the end caps.
 20. The method of claim 1 furthercomprising: interleaving composite plies of the stringer preforms with ametallic structure of the end caps.